U.S. patent number 10,286,577 [Application Number 15/144,808] was granted by the patent office on 2019-05-14 for composite mandrel for autoclave curing applications.
This patent grant is currently assigned to The Boeing Company. The grantee listed for this patent is The Boeing Company. Invention is credited to Panagiotis E. George, Brian G. Robins, Daniel M. Rotter, Todd J. Washburn.
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United States Patent |
10,286,577 |
Robins , et al. |
May 14, 2019 |
Composite mandrel for autoclave curing applications
Abstract
A composite mandrel includes a generally elongated mandrel body
comprising a resilient mandrel core and an elastomeric mandrel
outer layer disposed outside the mandrel core. A method for
fabricating a contoured stiffened composite panel is also
disclosed.
Inventors: |
Robins; Brian G. (Renton,
WA), Rotter; Daniel M. (Lake Forest Park, WA), Washburn;
Todd J. (Maple Valley, WA), George; Panagiotis E. (Lake
Tapps, WA) |
Applicant: |
Name |
City |
State |
Country |
Type |
The Boeing Company |
Chicago |
IL |
US |
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Assignee: |
The Boeing Company (Chicago,
IL)
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Family
ID: |
41504280 |
Appl.
No.: |
15/144,808 |
Filed: |
May 2, 2016 |
Prior Publication Data
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Document
Identifier |
Publication Date |
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US 20160243730 A1 |
Aug 25, 2016 |
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Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
Issue Date |
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12170843 |
Jul 10, 2008 |
9327467 |
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Current U.S.
Class: |
1/1 |
Current CPC
Class: |
B29D
99/0014 (20130101); B29C 33/3814 (20130101); B29C
70/865 (20130101); B29C 33/505 (20130101); B29L
2031/757 (20130101); B29L 2031/3076 (20130101) |
Current International
Class: |
B29C
33/38 (20060101); B29C 33/50 (20060101); B29C
70/86 (20060101); B29D 99/00 (20100101) |
References Cited
[Referenced By]
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Foreign Patent Documents
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JP |
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WO9851481 |
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Nov 1998 |
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WO |
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WO2005105402 |
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Nov 2005 |
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WO |
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WO2008003715 |
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Jan 2008 |
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WO |
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WO2008003721 |
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Jan 2008 |
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WO |
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WO2008003733 |
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Jan 2008 |
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WO |
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WO2010005811 |
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Jan 2010 |
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WO |
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Other References
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Application No. PCT/US2009/048889 (WO2010005811) 3 pages. cited by
applicant .
"Unitary", Webster's New Collegiate Dictionary, G. & C. Merriam
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Office Action, dated Sep. 27, 2010, regarding U.S. Appl. No.
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Final Office Action, dated Feb. 8, 2011, regarding U.S. Appl. No.
12/170,843, 11 pages. cited by applicant .
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Final Office Action, dated Apr. 25, 2012, regarding U.S. Appl. No.
12/170,843, 28 pages. cited by applicant .
Office Action, dated May 1, 2014, regarding U.S. Appl. No.
12/170,843, 18 pages. cited by applicant .
Final Office Action, dated Oct. 10, 2014, regarding U.S. Appl. No.
12/170,843, 13 pages. cited by applicant .
Office Action, dated Sep. 16, 2015, regarding U.S. Appl. No.
12/170,843, 10 pages. cited by applicant .
Office Action, dated Dec. 30, 2015, regarding U.S. Appl. No.
12/170,843, 15 pages. cited by applicant .
Office Action, dated Jun. 30, 2011, regarding U.S. Appl. No.
12/350,928, 16 pages. cited by applicant .
Final Office Action, dated Dec. 20, 2011, regarding U.S. Appl. No.
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Final Office Action, dated Aug. 7, 2012, regarding U.S. Appl. No.
12/350,928, 28 pages. cited by applicant .
Office Action, dated May 16, 2013, regarding U.S. Appl. No.
12/350,928, 13 pages. cited by applicant .
Office Action, dated Nov. 3, 2014, regarding U.S. Appl. No.
12/350,928, 17 pages. cited by applicant .
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cited by applicant.
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Primary Examiner: Gupta; Yogendra N
Assistant Examiner: Luk; Emmanuel S
Attorney, Agent or Firm: Yee & Associates, P.C.
Parent Case Text
This application is a divisional application of U.S. Pat. No.
9,327,467, filed Jul. 10, 2008, and issued May 3, 2016.
Claims
What is claimed is:
1. A composite mandrel, comprising: a generally elongated mandrel
body comprising: a resilient mandrel core; and an elastomeric
mandrel outer layer disposed outside the resilient mandrel core,
wherein a cross-sectional area and type of foam used for the
resilient mandrel core is engineered to impart compression
compliance under autoclave pressure to offset a combined thermal
expansion behavior of the resilient mandrel core and the
elastomeric mandrel outer layer.
2. The composite mandrel of claim 1, wherein the resilient mandrel
core comprises foam.
3. The composite mandrel of claim 1 wherein the generally elongated
mandrel body has a generally triangular cross-section.
4. The composite mandrel of claim 3 wherein the resilient mandrel
core comprises a core base, a pair of core sides extending from
said core base and a core apex extending between said pair of core
sides.
5. The composite mandrel of claim 4 wherein the core apex of the
resilient mandrel core is rounded.
6. The composite mandrel of claim 4 wherein the elastomeric mandrel
outer layer comprises a mandrel base disposed adjacent to the core
base of the resilient mandrel core; a pair of mandrel sides
disposed adjacent to the pair of core sides, respectively, of the
resilient mandrel core; and a mandrel apex disposed adjacent to the
core apex of the resilient mandrel core.
7. The composite mandrel of claim 6 wherein the mandrel apex of the
elastomeric mandrel outer layer is rounded.
8. The composite mandrel of claim 1, wherein the generally
elongated mandrel body has a trapezoidal cross-section.
9. The composite mandrel of claim 1, wherein the elastomeric
mandrel outer layer has a substantially constant thickness, and
wherein the elastomeric mandrel outer layer is configured to expand
uniformly during curing.
10. The composite mandrel of claim 1, wherein the elastomeric
mandrel outer layer is configured to be deformed for removal from a
composite structure.
11. The composite mandrel of claim 1, wherein the elastomeric
mandrel outer layer conforms to pad-ups and ramps.
12. The composite mandrel of claim 1, wherein the elastomeric
mandrel outer layer is in contact with and substantially
co-extensive with the resilient mandrel core.
13. A composite mandrel for fabricating an aircraft part,
comprising: a generally elongated one-piece mandrel body configured
to provide structural support to a cavity of the aircraft part
during a curing step in an autoclave and configured to be extracted
from the cavity after the curing step, the generally elongated
one-piece mandrel body comprising: a resilient mandrel core; and an
elastomeric mandrel outer layer disposed outside and in contact
with the resilient mandrel core.
14. The composite mandrel of claim 13 wherein the resilient mandrel
core comprises foam.
15. The composite mandrel of claim 13, wherein the generally
elongated one-piece mandrel body has a generally trapezoidal
cross-section, wherein the resilient mandrel core comprises a core
base, a pair of core sides extending from the core base and a
generally planar core top extending between the pair of core
sides.
16. The composite mandrel of claim 15 wherein the elastomeric
mandrel outer layer comprises a mandrel base disposed adjacent to
the core base of the resilient mandrel core; a pair of mandrel
sides disposed adjacent to the pair of core sides, respectively, of
the resilient mandrel core; and a mandrel top surface disposed
adjacent to the core top of the resilient mandrel core.
17. The composite mandrel of claim 13, wherein the elastomeric
mandrel outer layer is configured to be deformed for removal from
the cavity.
18. The composite mandrel of claim 13, wherein a cross-sectional
area and type of foam used for the resilient mandrel core is
engineered to impart compression compliance under autoclave
pressure to offset a combined thermal expansion behavior of the
resilient mandrel core and the elastomeric mandrel outer layer.
19. A composite mandrel for fabricating an aircraft part,
comprising: a generally elongated mandrel body configured to
provide structural support to a cavity of the aircraft part during
a curing step in an autoclave and configured to be extracted from
the cavity after the curing step, the generally elongated mandrel
body having a generally trapezoidal cross-section and comprising: a
resilient foam mandrel core having a core base, a pair of core
sides extending from said core base and a generally planar core top
extending between said pair of core sides; and an elastic rubber
mandrel outer layer disposed outside and in contact with the
resilient foam mandrel core and having a mandrel base disposed
adjacent to said core base of the resilient foam mandrel core; a
pair of mandrel sides disposed adjacent to said core sides,
respectively, of the resilient foam mandrel core; and a mandrel top
surface disposed adjacent to said core top of the resilient foam
mandrel core.
Description
TECHNICAL FIELD OF THE INVENTION
The disclosure relates to mandrels for forming cavities in
composite materials. More particularly, the disclosure relates to a
composite mandrel which is suitable for autoclave curing
applications in the formation of cavities in composite
materials.
BACKGROUND OF THE INVENTION
When composite materials are molded into shapes with cavities, such
as hat stringers, for example, there may be a need for some type of
tooling that can apply pressure from the cavity outward during the
curing step and can be extracted from the cavity after curing. The
existing tooling used for this purpose may include without
limitation inflatable rubber mandrels; solid mandrels such as
metal, rubber or composite mandrels; or dissolvable mandrels.
However, the inflatable rubber mandrels may be prone to leaking,
which may lead to widespread porosity in the resulting composite
laminate. The solid rubber mandrel may result in a cavity with a
distorted cross-sectional shape or exert an uneven pressure on the
composite laminate and may be too heavy for fabrication of large
parts. The solid metal or composite mandrels may not have
sufficient flexibility to be removed from parts having any degree
of curvature or complexity. The dissolvable mandrels may be
expensive to make and difficult to remove from large parts.
Existing mandrel designs may not accommodate the dimensional
changes of the composite part which occurs during application of
heat to the surrounding tooling and part materials at the curing
step. This can cause undesirable part material movement resulting
in such distortions as waviness, wrinkling and/or bridging in the
composite material.
Therefore, a mandrel is needed which is suitable for curing
applications in the formation of cavities in composite materials
and overcomes some or all of the limitations of conventional
composite mandrels.
SUMMARY OF THE INVENTION
The disclosure is generally directed to a composite mandrel. An
illustrative embodiment of the composite mandrel includes a
generally elongated mandrel body comprising a resilient mandrel
core and an elastomeric mandrel outer layer disposed outside the
mandrel core. The mandrel may combine the desired characteristics
of foam and rubber to produce a manufacturing aid for airplane
stringers or other similar open cavity parts made from fiber/resin
composite materials. The manufacturing aid which is embodied in the
composite mandrel may be less costly, more durable and less prone
to failures than current inflatable bladder technologies.
BRIEF DESCRIPTION OF THE ILLUSTRATIONS
FIG. 1 is a top view of an illustrative embodiment of the composite
mandrel.
FIG. 2 is a cross-sectional view, taken along section lines 2-2 in
FIG. 1, of the composite mandrel.
FIG. 3 is a cross-sectional view of an alternative illustrative
embodiment of the composite mandrel.
FIG. 4 is an exploded top view of a composite assembly, more
particularly illustrating insertion of multiple composite mandrels
into respective stiffening elements in the composite assembly
preparatory to curing of the composite assembly.
FIG. 5 is a cross-sectional view, taken along section lines 5-5 in
FIG. 4, of the composite assembly.
FIG. 6 is a top view of the composite assembly, with the composite
mandrels inserted in the respective stiffening elements of the
assembly.
FIG. 7 is a top view of the composite assembly, contained in vacuum
bagging preparatory to curing of the assembly.
FIG. 8 is an exploded top view of the composite assembly, more
particularly illustrating removal of the composite mandrels from
the respective stiffening elements in the composite assembly after
curing of the composite assembly.
FIG. 9 is a flow diagram which illustrates an illustrative method
for fabricating a contoured stiffened composite panel.
FIG. 10 is a flow diagram of an aircraft production and service
methodology.
FIG. 11 is a block diagram of an aircraft.
DETAILED DESCRIPTION
Referring initially to FIGS. 1 and 2, an illustrative embodiment of
the composite mandrel is generally indicated by reference numeral
1. The composite mandrel 1 may be used to fill a cavity (not shown)
in an airplane stringer or other open-cavity part (not shown) made
from fiber/resin composite materials to prevent collapse of the
cavity during curing of the composite materials. The composite
mandrel 1 may be less costly, more durable and more effective and
reliable than current inflatable bladder mandrel technologies.
The composite mandrel 1 includes a generally elongated mandrel body
7 having a mandrel core 2 which is a resilient material and a
mandrel outer layer 10 which is disposed outside the mandrel core
2, as shown in FIG. 2, and is an elastomeric material. In some
embodiments, the mandrel core 2 is foam or other such material
which incorporates open space and/or air pockets to prevent bulk
modulus behavior during thermal expansion and the mandrel outer
layer 10 may be an elastomeric material such as elastic rubber, for
example and without limitation. The mandrel core 2 and the mandrel
outer layer 10 may be generally coextensive with the mandrel body
7.
The mandrel core 2 and the mandrel outer layer 10 may have any
cross-sectional shape depending on the particular use requirements
of the composite mandrel 1. In some applications, for example, each
of multiple composite mandrels 1 may be suitably configured to fill
respective stiffening elements (such as stringers) 27 during the
curing and/or cocuring of a composite panel assembly 24, as shown
in FIGS. 4-8 and will be hereinafter described. As shown in FIG. 2,
in some embodiments of the composite mandrel 1, the mandrel body 7
may have a generally triangular cross-sectional shape. Accordingly,
the mandrel core 2 has a generally flat or planar core base 3 with
lateral core edges 6. Core sides 4 angle from the respective core
edges 6. A core apex 5, which may be rounded, extends between the
core sides 4. The shape of the mandrel outer layer 10 may generally
correspond to that of the mandrel core 2, defining a mandrel base
11 which extends adjacent to the core base 3; a pair of mandrel
sides 12 which extend adjacent to the respective core sides 4; a
mandrel apex 13 which may be rounded and is disposed adjacent to
the core apex 5; and mandrel edges 14 which correspond positionally
to the respective core edges 6 of the mandrel core 2.
As shown in FIG. 3, in some embodiments of the composite mandrel
1a, the mandrel body 7a may have a generally trapezoidal shape.
Accordingly, the mandrel core 2a has a generally flat or planar
core base 3; a pair of core sides 4 which angle from the core base
3; and a generally flat or planar mandrel core top 8 which extends
between the core sides 4. The mandrel outer layer 10a defines a
mandrel base 11 which extends adjacent to the core base 3; a pair
of mandrel sides 12 which extend adjacent to the respective core
sides 4; a generally flat or planar mandrel top surface 16 which is
disposed adjacent to the mandrel core top 8; and mandrel edges 14
which correspond to the respective core edges 6 of the mandrel core
2a.
Referring next to FIGS. 4-8, in typical application, multiple
composite mandrels 1 are inserted in respective stiffening elements
27 provided in a stiffening layer 26 of a composite panel assembly
24 during curing of the composite panel assembly 24. The composite
panel assembly 24 will ultimately form an airplane stringer (not
shown); however, it will be appreciated by those skilled in the art
that the composite mandrels 1 can be adapted to fill cavities in
any other type of open-cavity or closed-cavity composite material
part made from fiber/resin composite materials during curing of the
composite material part. The composite mandrels 1 can be adapted to
fill cavities having a constant cross-sectional shape or a
cross-sectional shape which varies along the length of the
composite material, such as cavities which taper or curve along the
length of the cavity, for example and without limitation.
As illustrated in FIG. 5, in an embodiment of fabrication of the
composite panel assembly 24, a base composite layer 25 may
initially be placed on a tooling surface 20 of OML tooling or IML
tooling, for example and without limitation. The tooling surface 20
may have a generally concave contour, as shown. Alternatively, the
tooling surface 20 may have a generally planar or convex contour,
depending on the particular application. The stiffening layer 26
may be placed on the base composite layer 25. The stiffening
elements 27 may be shaped in the stiffening layer 26 and extend
along the longitudinal axis of the tooling surface 20 in generally
parallel relationship with respect to each other, as shown in FIG.
4, and in generally perpendicular relationship with respect to the
concave contour of the tooling surface 20. Alternatively, the
stiffening elements 27 may be separate or discrete units. As
further shown in FIG. 5, each stiffening element 27 has a
stiffening element cavity 28. In some embodiments, the stiffening
elements 27 may be oriented in orientations other than along the
longitudinal axis of the tooling surface 20 and may converge or
diverge, for example and without limitation.
As shown in FIGS. 4 and 6, multiple composite mandrels 1 may be
inserted into the stiffening element cavitys 28 of the respective
stiffening elements 27. The elastomeric mandrel outer layer 10 of
each composite mandrel 1 allows for a proper fit of the composite
mandrel 1 into the stiffening element cavity 28 of each stiffening
element 27 and conforms to pad-ups and ramps. As shown in FIG. 7,
the composite panel assembly 24 may then be enclosed in vacuum
bagging 30 and cured by autoclaving. During the curing process, the
composite mandrels 1 maintain the shape and prevent collapse of the
respective stiffening elements 27 as the composite material of the
base composite layer 25 and the stiffening layer 26 hardens.
After curing, the composite panel assembly 24 is removed from the
vacuum bagging 30. The composite mandrels 1 may be removed from the
stiffening element cavitys 28 of the respective stiffening elements
27, as shown in FIG. 8. During removal, the elastomeric mandrel
outer layer 10 of each composite mandrel 1 may easily be deformed;
this reduces the effort required for removal. The cured composite
panel assembly 24 may then be processed to complete fabrication of
the airplane assembly (not shown) or other composite part,
according to the knowledge of those skilled in the art.
It will be appreciated by those skilled in the art that the
resilient mandrel core 2 of the composite mandrel 1 enhances the
structural and compressive characteristics of the composite mandrel
1 relative to the designs of conventional mandrels. This structural
and compressive support may be necessary to maintain the shape of
the stringer or other composite part during automated composite
fiber placement as well as autoclave curing. Since the outer
mandrel layer 10 may be a constant thickness, it may expand
uniformly during curing, thus avoiding the problems associated with
uneven expansion of a solid rubber material. The cross-sectional
area and type of foam used for the mandrel core 2 may be engineered
to impart compression compliance under autoclave pressure, thus
offsetting the combined thermal expansion behavior of the foam and
rubber.
Referring next to FIG. 9 of the drawings, a flow diagram 900 which
illustrates an illustrative method for fabricating a contoured
stiffened composite panel is shown. In block 902, a tooling
surface, such as the tooling surface 20 which was heretofore
described with respect to FIG. 5, for example and without
limitation, is provided. The tooling surface may have a concave,
planar, convex or alternative contour. In block 904, a base
composite layer is laminated on the tooling surface. In block 906,
open-section stiffening elements are positioned on the base
composite layer. In block 908, composite mandrels are provided.
Each composite mandrel includes a resilient mandrel core and an
elastomeric mandrel outer layer disposed outside the resilient
mandrel core. In block 910, composite mandrels are inserted in the
respective stiffening elements. In block 912, the composite panel
and stiffening elements are sealed in vacuum bagging. In block 914,
the composite panel and the stiffening elements are cured. An
autoclave may be used during curing. In block 916, the composite
mandrels are removed from the stiffening elements.
Referring next to FIGS. 10 and 11, embodiments of the disclosure
may be used in the context of an aircraft manufacturing and service
method 78 as shown in FIG. 10 and an aircraft 94 as shown in FIG.
11. During pre-production, exemplary method 78 may include
specification and design 80 of the aircraft 94 and material
procurement 82. During production, component and subassembly
manufacturing 84 and system integration 86 of the aircraft 94 takes
place. Thereafter, the aircraft 94 may go through certification and
delivery 88 in order to be placed in service 90. While in service
by a customer, the aircraft 94 may be scheduled for routine
maintenance and service 92 (which may also include modification,
reconfiguration, refurbishment, and so on).
Each of the processes of method 78 may be performed or carried out
by a system integrator, a third party, and/or an operator (e.g., a
customer). For the purposes of this description, a system
integrator may include without limitation any number of aircraft
manufacturers and major-system subcontractors; a third party may
include without limitation any number of vendors, subcontractors,
and suppliers; and an operator may be an airline, leasing company,
military entity, service organization, and so on.
As shown in FIG. 11, the aircraft 94 produced by exemplary method
78 may include an airframe 98 with a plurality of systems 96 and an
interior 100. Examples of high-level systems 96 include one or more
of a propulsion system 102, an electrical system 104, a hydraulic
system 106, and an environmental system 108. Any number of other
systems may be included. Although an aerospace example is shown,
the principles of the invention may be applied to other industries,
such as the automotive industry.
The apparatus embodied herein may be employed during any one or
more of the stages of the production and service method 78. For
example, components or subassemblies corresponding to production
process 84 may be fabricated or manufactured in a manner similar to
components or subassemblies produced while the aircraft 94 is in
service. Also, one or more apparatus embodiments may be utilized
during the production stages 84 and 86, for example, by
substantially expediting assembly of or reducing the cost of an
aircraft 94. Similarly, one or more apparatus embodiments may be
utilized while the aircraft 94 is in service, for example and
without limitation, to maintenance and service 92.
Although the embodiments of this disclosure have been described
with respect to certain exemplary embodiments, it is to be
understood that the specific embodiments are for purposes of
illustration and not limitation, as other variations will occur to
those of skill in the art.
* * * * *